The airfoils of gas turbines are typically made of nickel or cobalt based superalloys which show high resistance against the hot and corrosive combustion gases present in gas turbine. However, although such superalloys have considerably high corrosion and oxidation resistance, the high temperatures of the combustion gases in gas turbines require measures to improve corrosion and/or oxidation resistance further. Therefore, airfoils of gas turbine blades and vanes are typically at least partially coated with a thermal barrier coating system to prolong the resistance against the hot and corrosive environment. In addition, airfoil bodies are typically hollow so as to allow a cooling fluid, typically bleed air from the compressor, to flow through the airfoil. Cooling holes present in the walls of the airfoil bodies allow a certain amount of cooling air to exit the internal passages so as to form a cooling film over the airfoil surface which further protects the superalloy material and the coating applied thereon from the hot and corrosive environment. In particular, cooling holes are present at the trailing edges of the airfoils as it is shown in U.S. Pat. No. 6,077,036, U.S. Pat. No. 6,126,400, US 2009/0194356 A1 and WO 98/10174, for example.
Trailing edge losses are a significant fraction of the over all losses of a turbomachinery blading. In particular, thick trailing edges result in higher losses. For this reason, cooled airfoils with a cutback design at the trailing edge have been developed. This design is realised by taking away material on the pressure side of the airfoil from the trailing edge up to several millimeters towards the leading edge. This measure provides very thin trailing edges which can provide big improvements on the blading efficiency. An airfoil with a cutback design and a thermal barrier coating is, for example, disclosed in WO 98/10174 A1. However, the beneficial effect on the efficiency can only be achieved if the thickness of the trailing edge is rather small. On the other hand for a blade with thermal barrier coating, the combined thickness of the cast airfoil body wall and the applied thermal barrier coating system exceeds the optimum thickness of the design. In addition, as the flow velocity of the gas is the greatest at the trailing edge of the airfoil a thermal barrier coating applied to the trailing edge is prone to high levels of erosion.
It is known to selectively provide a thermal barrier coating system to the airfoil, in particular such that the trailing edge of an airfoil and adjacent regions of an airfoil remain uncoated. Selective coatings are, for example, described in U.S. Pat. No. 6,126,400, U.S. Pat. No. 6,077,036 and, with respect to the coating method, in US 2009/0104356 A1.
However, in U.S. Pat. No. 6,077,036 the pressure side of the airfoil is completely uncoated which means that areas which would not suffer from a higher combined thickness of the cast airfoil body and the coating applied thereon remain unprotected against the temperature the hot combustion gas.
WO 2008/043340 A1 describes a turbine airfoil with a thermal barrier coating the thickness of which varies over the airfoil surface. However, like in WO 98/10174 the trailing edge is fully coated so that the beneficial effect on blading efficiency can not be achieved. In U.S. Pat. No. 6,126,400 the thermal barrier coating only covers about half the airfoil, as seen from the leading edge towards the trailing edge.
In US 2009/0104356 A1 the method of masking the trailing edge will produce a step in the coating which adversely affects the aerodynamics of the blade.